1. Field of the Invention
The invention relates generally to rocket engines and, in particular, to an aerospike rocket engine.
2. Background Art
Conventional rocket engines use round, bell-shaped nozzles. These nozzles, however, have an inherent limitation in that the combustion gas, or plume inside the nozzle can expand only as far as the shape and length of the nozzle allow, resulting in substantial under and/or over expansion, with a resulting loss of thrust and instability/vibration of the expanding plume. Bell nozzles are, therefore, typically designed for specific applications, e.g., take-off, high altitude, or outer space. However, even within the confines of these applications, under/over expansion invariably occurs due to 1) changes in atmospheric pressure, and 2) a finite expansion capability of approximately 1:400 (where infinite expansion is theoretically required in space), which may result in up to 5% loss of thrust. See, e.g., Missile Engineering Handbook, van Nostrand, FIG. 7.1.1, 1957; Aviation Week and Space Technology, p. 130, Aug. 10, 1987). Therefore, a bell nozzle having a given size and shape can reach peak efficiency only at an altitude where the plume expansion within the nozzle equals the theoretical expansion that would be permitted by the atmospheric pressure at that altitude.
To overcome the bell nozzle""s limitation, Rocketdyne Propulsion and Power (xe2x80x9cRocketdynexe2x80x9d), a subsidiary of the Boeing Co., developed a nozzle which resembles a bell nozzle turned inside-out called an xe2x80x9caerospikexe2x80x9d nozzle. More specifically, a linearized version of the aerospike nozzle called a xe2x80x9clinearxe2x80x9d aerospike nozzle was developed for the proposed X33/VentureStar single-stage-to-orbit (xe2x80x9cSSTOxe2x80x9d) space plane project. The linear aerospike engine resembles a bell-shaped nozzle that has been split in half and the two halves put back-to-back to each other, and the end of nozzle clipped or truncated. In some cases, however, the linear aerospike engine may have only one of the two halves, i.e., a single-sided engine. Because the plume of the aerospike nozzle is manifested on the peripheral of the nozzle, it is free to expand, limited only by atmospheric pressure. As a rocket using the aerospike nozzle climbs higher and higher, the plume is able to expand continuously against the decreasing atmospheric pressure, albeit at a cost to the thrust vector which diverges progressively sideways.
Referring to FIG. 1, there is shown a bank of five linear aerospike engines 10 arranged side-by-side. Each aerospike engine 10 comprises a rectangular wedge or tapered body 12, a slanted or curved reaction surface or plane 14, a leading end 16 and a trailing end 18. Each engine 10 has at least one injector 20 or, more typically, a set of injectors 20 adjacent the leading end 16 and arranged to direct a propellant or fuel down the reaction plane 14 towards the trailing end 18. Upon combustion of the propellant or fuel from the injector 20, the combustion gas, or plume, travels down the reaction plane 14 and exerts propulsive pressure on the reaction plane 14, which provides the thrust for the space plane.
As can be seen, turning to FIGS. 2A-2C, the linear aerospike design allows the plume to expand freely against atmospheric pressure. At low altitude, the exhaust plume 24 is held in a fairly narrow band 26 by the high atmospheric pressure as shown in FIG. 2A. However, referring to FIG. 2B, at high altitude and low atmospheric pressure, the plume 24 is able to expand. Shock waves produced by the supersonic speed of the space plane at high altitude provides a shock front 28 that can assists in resisting the expansion of the plume 24. As the space plane 22 climbs into outer space, the vacuum of space may tend to pull the plume 24 away from the reaction plane 14, as shown in FIG. 2C. This may result in xe2x80x9cdivergence,xe2x80x9d wherein the plume""s 24 thrust vectors becomes misaligned with the direction of flight, resulting in a decrease in net thrust and, hence, engine efficiency.
One solution to this divergence syndrome is to extend the reaction plane 14 so as to facilitate full expansion of the plume 24. However, because the plume 24 is unconfined, the boundary layer may tend to separate from the reaction plane 14. Boundary layer separation is a lifting off or peeling away of the plume 24 from the reaction plane 14. According to Bernoulli""s law, as long as the boundary layer remains sufficiently energized, the plume 24 will adhere to the reaction plane 14 by virtue of the negative pressure between the high-speed boundary layer and the reaction plane 14. As the plume 24 travels along the reaction plane 14, the boundary layer may run out of energy and separate from the reaction plane 14. The effects of boundary layer separation include instability or turbulence which can produce severe mechanical vibrations that can damage the space plane 22. In addition, boundary layer separation may result in a loss of thrust and engine efficiency. Separation usually starts at the end of the boundary layer where the energy of the boundary layer is low. Atmospheric pressure can help to hold the plume 24 against the reaction plane 14. Therefore, separation is more likely to occur at high altitude where the atmospheric pressure is low.
One way of preventing boundary layer separation is by truncating the reaction plane 14 so that the reaction plane 14 is shorter (as can be seen in published illustrations of the X33). This allows the boundary layer to traverse the entire length of the reaction plane 14 before running out of energy. The trade-off, however, is that there is a reduction in thrust and engine efficiency relative to an untruncated reaction plane due to 1) under expansion, and 2) thrust vector diversion/deflection. Furthermore, the shorter reaction plane 14 may not allow the propellant or fuel sufficient time to completely combust/accelerate before reaching the end of the reaction plane 14, which can result in reduced thrust on the reaction plane 14. This reduction in thrust may be critical at high altitudes where the space plane needs to attain very high velocity.
Over and above the truncation limitation of the X33 implementation of the aerospike engine, scaling up of the aerospike plan form to suit larger space plane applications (e.g., the proposed VentureStar heavy lift shuttle) may additionally require cascaded or staged propellant/fuel injection in lieu of the impact of dimensional scaling.
As mentioned above, a conventional linear aerospike engine may be inefficient for powering very large space planes or other vehicles because of the reduction or loss of pressure due to truncation of the engine wedge. The present invention provides means for maintaining and/or increasing the pressure across the reaction plane to thereby enhance the thrust of the engine, and for reducing the divergence or deflection of the thrust vectors. The present invention also provides means for preventing or inhibiting boundary layer separation from the reaction plane.
In general, in one aspect, the invention is related to a rocket engine comprising a tapered body, a slanted reaction plane on the body, and means for increasing propulsive pressure on the reaction plane. In one embodiment, the means for increasing propulsive pressure on the reaction plane may be a first fuel injector adjacent a leading end of the engine and injecting a first fuel on the reaction plane and a second fuel injector between the leading end and a trailing end of the engine and injecting a second fuel on the reaction plane. The first fuel and the second fuel may be cascaded on the reaction plane, and may be of the same type, or two different types of fuels.
In another embodiment, the means for increasing propulsive pressure on the reaction plane may be a means for inducing a vortex on the reaction plane substantially parallel to a lateral axis of the reaction plane. The induced vortex may draw ambient air towards the reaction plane. The vortex may be induced by at least two fuel injectors configured to inject fuel on the reaction plane in counter current directions, an adjustable spoiler adjacent a leading end of the engine, or a rotatable turbine spaced apart from and perpendicular to the reaction plane adjacent a leading end of the engine.
In still another embodiment, the means for increasing propulsive pressure on the reaction plane may be a reaction plane extension. The reaction plane extension may be selectively retractable into a trailing end of the engine, or it may be selectively foldable onto a trailing end of the engine.
In general, in another aspect, the invention relates to a rocket engine comprising
a tapered body, a slanted reaction plane on the body, and means for inhibiting boundary layer separation from the reaction plane. In one embodiment, the means for inhibiting boundary layer separation may be a rotatable turbine. perpendicular to and spaced apart from the reaction plane adjacent a leading end of the engine. The turbine may have circumferential grooves and/or strings of Wheeler vortex generators on the outer circumference of the turbine, or the turbine may have dimples or perforations on the outer circumference of the turbine.
In another embodiment, the means for inhibiting boundary layer separation be a means for inducing vorticity on the reaction plane. The means for inducing vorticity may include Wheeler vortex generators attached to or otherwise formed on a leading end of the engine and/or the reaction plane.
In still another embodiment, the means for inhibiting boundary layer separation may be a coating of a special drag-resistant material, e.g., Teflon on the reaction plane. The means may also be longitudinal grooves, or dimples on the reaction plane.
In yet another embodiment, the means for inhibiting boundary layer separation includes a reaction plane extension having slits and/or perforations thereon. The extension may include covers attached to the extension adjacent the slits for selectively covering the slits. The extension may be selectively retractable into a leading end of the engine, or selectively foldable onto a leading end of the engine. The trailing end of the extension may have slits thereon and covers adjacent the slits for selectively covering the slits.
In yet another embodiment, there may be a partition attached perpendicular to the reaction plane and extending parallel to a longitudinal axis of the reaction plane.
In general, in another aspect, the invention relates to a rocket engine comprising
a tapered body, a slanted reaction plane on the body, means for increasing propulsive pressure on the reaction plane, and means for inhibiting boundary layer separation from the reaction plane.
In general, in another aspect, the invention relates to a space plane comprising
a main body, an aerospike engine attached to the main body, and a bell-shaped nozzle engine attached to the main body. In one embodiment, the aerospike engine includes a tapered body and a slanted reaction plane on the tapered body, and further comprises means for increasing propulsive pressure on the reaction plane and means for inhibiting boundary layer separation.
In general, in another aspect, the invention relates to a space plane comprising
a first aerospike engine having first tapered body and first slanted reaction plane on the first body, a second aerospike engine having second tapered body and second slanted reaction plane on the second body, wherein the first and second, reaction planes are of different lengths. In one embodiment, the first reaction plane is shorter than the second reaction plane. In another embodiment, the first engine uses a different type of fuel than the second engine. The space plane may also include means for increasing propulsive pressure on the first and second reaction planes, and means for inhibiting boundary layer separation.
In general, in another aspect, the invention relates to a method of operating a linear aerospike engine having a tapered engine body which has a slanted reaction plane, wherein the method comprises injecting a first fuel towards the reaction plane, and injecting a second fuel towards the reaction plane, wherein the first fuel and the second fuel are cascaded on the reaction plane. In one embodiment, the method comprises using a first fuel which may be substantially the same as the second fuel, or using a first fuel which may be a different type than the second fuel. In another embodiment, the method comprises selectively modulating the injection of the first and second fuel in accordance with a predetermined injection strategy.
In general, in another aspect, the invention relates to a method of operating a linear aerospike engine having a tapered engine body which has a slanted reaction plane, wherein the method comprises firing the engine, and inducing a vortex substantially parallel to a lateral axis of the reaction plane. In one embodiment, the vortex may be induced using an adjustable spoiler adjacent to a leading end of the reaction plane, a rotatable turbine perpendicular to and spaced apart from the reaction plane adjacent a leading end of the engine, or at least two fuel injectors configured to direct fuel in counter current directions on the reaction plane.
In general, in another aspect, the invention relates to a method of operating a space vehicle having a tapered engine body which has a slanted reaction plane, wherein the method comprises firing the engine, extending the reaction plane at a predetermined time, and removing decelerated boundary layer fluid from a boundary layer while the reaction plane is extended. In one embodiment, the decelerated boundary layer fluid is removed using slits and/or perforations in an extended portion of the reaction plane.
In general, in another aspect, the invention relates to a method of operating a space vehicle having first and second linear aerospike engines, wherein the first engine has a different length reaction plane than the second engine, the method comprising firing the first and second engines essentially at the same time. In one embodiment, the first engine may have a shorter reaction plane than the second engine. In another embodiment, the length of the second engine""s reaction plane may be extended at a predetermined time interval. In another embodiment, the method further comprises selectively adapting the reaction plane length of the first and second engine while the space vehicle is in flight. In yet another embodiment, the thrusts of the first and second engines may be selectively adapted while the space plane is in flight. In yet another embodiment, the first and second engines are selectively modulated in accordance with a predetermined firing strategy. In yet another embodiment, the space plane includes a bell-shaped nozzle engine which may be fired at essentially the same time as the first and second engines, or after the space plane reaches outer space.
Other aspects and advantages of the invention will be apparent from the following description and the appended claims.